The invention is a minimally intrusive slender supersonic injector flush mounted to the wall of a combustor through which combustion air flows. When not in use, the invention is substantially nonintrusive. The invention may also be used to control the attitude of a spaceplane above the atmosphere of the earth.
A typical scramjet engine includes a combustor having a chamber wherein a fuel-air mixture moving at supersonic speed is burned. At least one fuel injector directs supersonically-moving fuel such as pressurized hydrogen into the chamber. The engine also includes an air inlet, which delivers compressed supersonically-moving air to the combustor chamber and further includes an exhaust nozzle which channels the burning gas out of the combustor chamber to help produce the engine thrust. The fuel injector discharge orifices are the openings in the combustor chamber to which fuel is delivered by a fuel system which may includes tanks, pumps and conduits.
State-of-the-art flush mounted circular and wedge injectors, similar to those proposed for use in scramjet (supersonic combustion ramjets) and rocket-based, combined-cycle (RBCC) air breathing space planes or missiles and liquid oxygen-augmented nuclear thermal rocket (LANTR) proposed Mars space transportation propulsion system, are intrusive when not used (generating significant shock disturbances throughout the system) and when used, create a high heat flux near the origin of injection due to a created separation bubble in front of the jet and/or recirculation eddies just downstream of the jet. Penetration and adequate mixing of exiting jet or jets in the main supersonic flow that supports a burnable/stable combustion process before being discharged is difficult to obtain.
RBCC/Scramjet/LANTR propulsion systems require fuel and/or oxidizer augmentation injectors that fulfill specific penetration, mixing, and uniform stable burning needs within the shortest distance for a wide range of supersonic cross flow Mach numbers when in use, but must be minimally intrusive when not in use. A flush mounted jet in a supersonic cross flow is quickly deflected by aerodynamic effects until the plume becomes parallel to the combustor surface freestream cross flow but does not mix adequately to support stable supersonic combustion. A plyup/ramping in the separated boundary layer occurs in front of the blockage produced by the state-of-the-art jet plume bringing the regional subsonic boundary layer flow nearly to rest. A recirculation pattern occurs withing this region enlarging the angular separated boundary layer plyup. A large bow pressure disturbance wave results as reported in NASA/TM-1999-208893 which is incorporated by reference hereto. This phenomena continues to persist affecting stable RBCC/scramjet/combustion performance to date regardless of the prior art injector angle and/or whether such injector is followed by a high drag mixing cavity.
U.S. Pat. No. 5,202,525 issued to Coffinberry on Apr. 13, 1993 states that the mixing of hydrogen fuel in a scramjet combustor is a difficult process since the compressed airfow is flowing at supersonic velocities with substantial momentum and the fuel injected into the combustor has relatively low momentum. Coffinberry further states that oxygen and nitrogen molecules contained in the air have relatively large mass inertia which typically easily overcome the relatively low mass inertia of molecular hydrogen in the fuel. Accordingly, hydrogen fuel has the tendency to simply follow the stream of supersonic airflow without significant mixing. In order to obtain acceptable combustion in the scramjet combustor, acceptable mixing of the fuel and supersonic airflow must be obtained. See, the ""525 patent, col. 1, lns. 29-42. Supersonic combustors face the further challenge that the fuel must be fully mixed within the combustor in a length as short as possible.
U.S. Pat. No. 5,280,705 issued Jan. 25, 1994 to Epstein et al. discloses an intermittent admission of the fuel to the airflow to promote enhanced combustion and to minimize the heat load on the combustor.
U.S. Pat. No. 5,220,787 issued to Bulman Jun. 22, 1993 discloses a locally pressure matched injector. Namely, the exit pressure of the injector is matched to the cross-flow pressure of the combustion air. By matching the exit pressure of the fuel jet to the pressure surrounding the fuel jet, a jet of the narrowest width is produced having the highest momentum. Bulman cites F. S. Billig et al and other researchers (Billig, F. S., Orth, R. C., Lasky, M., xe2x80x9cA Unified Analysis of Gaseous Jet Penetration,xe2x80x9d American Institute of Aeronautics and Astronautics Journal, Vol. 9, No. 6, Jun. 1971, pp. 1048-1058, that studied the penetration and mixing of fuel jets in cross flows in the 1960""s. Billig et al. studied the effects of introducing fuel through both a circular opening and a noncircular opening. According to Bulman, the use of these noncircular openings by Billig did not have the desired effect, namely, improved penetration and mixing because the pressure matching was only performed on an average basis.
The Bulman ""787 patent cites the need for better fuel mixing which improves the combustion efficiency of the engine. Bulman recognized prior attempts to get more fuel in the air stream by simply pumping more fuel therein. However, according to Bulman, air/fuel mixing is not well served by having a few large injectors because the result is a large over-fueled region surrounded by underfueled air. See, for example, the Bulman ""787 patent at col. 1, lns. 37-45. Bulman cites the need for better mixing in relation to the gap between injectors. In other words, according to Bulman, better penetration and mixing enables the use of multiple injectors spaced closer together which promotes more thorough and consistent combustion in a smaller space within the engine.
Bulman ""787 discloses a fuel injector having at least one fuel inlet port, throat and fuel exit port serially connected and which in combination produce the local pressure match as well as a low drag shape. Bulman cites calculation of the throat contour to produce the correct area ratio to achieve the local pressure matching condition in that the fuel jet exiting from the fuel injector from the proximate end is lower in velocity and therefore at a higher pressure while the fuel jet exiting from the injector body near the distal end has a higher velocity and lower local pressure. See, col. 10, lns. 39-48, of the Bulman ""787 patent. The ratio of the exit width to the throat width determines the local area ratio. Bulman ""787 in FIG. 13A thereof cites an example having a 2.25 degree throat area half angle.
The Bulman ""787 seven (7) degree half angle wedge and its cascade injector version reduced forward hot spots somewhat and slightly increased penetration. However, as tested this injector consistently produced a non-uniform velocity profile jet (Mach 2.1 to 1.67) from bow to stern over a wide range of injectant pressures. State-of-the-art scramjet engines presently use groups of normal or angled hole injectors (quarter inch diameter) just forward of high drag/step mixing cavities or the Bulman ""787 wedge injectors positioned on a strut within the combustor duct geometry.
All prior art supersonic injectors are intrusive when not used, creating a not to be ignored pressure wave and shock disturbances that propagate downstream the combustor duct. When supersonic flush injection is used within or into supersonic crossflow a large pressure wave, shock disturbances and hot spots occur near the jet. See, NASA/TM 107533, NASA/TM 1999-208893, and/or NASA/TM-2001-210951.
State-of-the-art hole injectors exhibit a forward bow wave which tends to separate extensively forward of the bow shock and creates a large disturbance (circulations and hot spots) downstream. Bulman""s structure reduces the forward disturbance significantly but does not eliminate it. Downstream of the Bulman injector the disturbance is significant with eddies and potential hot spots existing, reference NASA/TM-2001-210951 FIGS. 5(a), 5(b) and 5(c), composite PLIF (PLANAR LASER-INDUCED FLUORESCENCE) plume images 1.50 and 2.00 inches downstream of the jets origin of injections.
The instant invention will be better understood when reference is made to the following Summary of the Invention, Brief Description of the Drawing, Description of the Invention and Claims.
The invention is a slender fuel injector comprising an elongated body flush mounted to the wall of a combustor. An opening having a substantially hour-glass shape in cross section extends from bow to stern of the injector. The substantially hour glass-shape in cross section may also be described as a venturi shape in cross section. Bow, front, fore and proximate are terms used to describe the portion of the injector which is first traversed by the incoming cross flow of combustion air. Stem, back, aft and distal are terms used to describe the portion of the injector which are last traversed by the incoming cross flow of combustion air.
Typical fuels include hydrogen, JP5, methane, propane, methylcyclohexane, pentaborane and mixtures of these. For typical oxidizer augmentation application, Liquid Oxygen Augmented Nuclear Thermal Rocket (LANTR), near space-air breathing supersonic/hypersonic combustion application, gaseous oxygen or liquid oxygen may be injected using the injectors of this invention.
The fuel injector of the present invention may be used in a combustor in different configurations, for example, in a convex surface, in a concave surface, or they may be curved (i.e., arced). Additionally, they may be curved and used in a convex and/or in a concave surface. Several injectors will be used in a combustor simultaneously. Straight line injectors can be angled or arced to starboard followed by downstream rows angled or arced to port such that penetration from the first row becomes further lifted by displacement by the second row while undergoing augmented mixing by the counter swirl patterns setup by the second row. In addition, the injectors may be flush mounted in a concave or convex surface.
The injectors may be used in other applications such as to control the attitude of a spaceplane operating above the atmosphere of the earth. In this application the injector may emit air or other compressed gas to control the orientation of the spaceplane. The injectors may also be used to supply oxidizers for combustion.
The opening in the elongated body of the injector includes a throat which diverges (increases) in width and depth from bow to stern. An hour glass shape (in cross section) opening diverges (increases) in width and depth from bow to stern. A starboard (right) and a port (left) block comprise the injector which is flush mounted to the combustor. Spacers fore and aft are employed to control the precise throat area of the injector. The throat is defined as the cross sectional area of the injector at the area of minimum cross section of the venturi or hour glass.
The surface of the injector is machined from 2124 Aluminum so as to obtain a 16 RMS (root mean square) surface roughness. The injector may be fabricated from various metals and/or ceramics. Direct metal deposition technology may be used to form the injectors. A one-ten thousandths stepover machining process is employed to obtain the 16 RMS surface roughness. Injectors of the present invention may be made from metal such as aluminum or stainless steel or they may be made from the investment casting process. Aluminum has better machinability than stainless steel. The injectors may be made from ceramics. Injectors may be made from an integral piece of material or they may be made in halves and used in conjunction with spacers. The injectors may also be made from the investment casting process in an integral piece.
The injectors of the invention are Mach number specific. A family of injectors may be used in a combustor to accommodate various conditions (altitude and velocity). In general, as the Mach number of the cross flow increases, the throat spacing from one side (starboard side) of the injector to the other side (port side) of the injector decreases and the length of the injector increases. In other words, a general characteristic of the injector is that it has a long and slender throat which is formed in an elongated body.
The invention creates only a minute disturbance when injecting fuel or oxidizer because it is slender and has a constant (uniform) injection velocity from bow to stern. The half angle of the throat area for the instant invention is 1.11 degrees for a Mach 2 injector and is 0.03 degrees for a Mach 4 injector. As the Mach number of the injector increases the half angle of the throat area decreases and the length of the injector increases. It is anticipated that the injectors of the instant invention will be used in air breathing flight vehicle applications that achieve speeds at least as high as Mach 25 to Mach 30. Compressed air velocity/flow within combustors for these applications will be high supersonic to low hypersonic (Mach 4.5 to Mach 7.0) speed.
The uniform velocity of the injectant penetrates farther into the supersonic mainstream cross flow with better dispersion and mixing beginning at the exit plane of the injector. The fuel injector includes a stern section which includes a stern wall. The stern wall is angled in the direction of the cross flow of the combustion air so as to substantially eliminate any disturbance of the injector when not in use. Preferrably, the stern wall is angled 5 to 10 degrees in the direction of the cross flow of the combustion air. A fan radius of approximately 0.0864 inches is present on the stern wall and the flush portion of the injector to allow a (Prandt-Meyer expansion) injection release.
The momentum flux ratio, J, is used to characterize the penetration of a jet into a cross flow. The momentum flux ratio, J, is defined as:   J  =                    γ        inj            ⁢              p        inj            ⁢              M        inj        2                            γ        tun            ⁢              p        tun            ⁢              M        tun        2            
where xcex3 is the ratio of specific heat for air, p is pressure and M is the injector Mach number, and the subscripts xe2x80x9ctunxe2x80x9d and xe2x80x9cinjxe2x80x9d refer to the tunnel flow and the injectant flow, respectively. While the momentum flux ratio is important and may be increased by raising the injection pressure and/or the Mach number, it has been discovered that it is very important to distribute the injectant with uniform velocity from the bow to the stern of the injector thereby penetrating farther (wider and higher) into the supersonic mainstream cross flow. In the injector of the instant invention J, the momentum flux ratio, is uniform from bow to stern. J is uniform from bow to stern for all of the injectors of the instant invention regardless of Mach number. Uniform velocity along the slender injector of the invention (as tested in the supersonic wind tunnel at the NASA Glenn Research facility in Cleveland, Ohio) resulted in better dispersion as compared to the prior art injector of Bulman (7 degree half angle wedge of the Bulman ""787 patent) with the same mass flow.
Substantial wind tunnel analyses have been performed on the instant invention, the prior art Bulman ""787 patent 7 degree half-angle wedge and the prior art (0.25 inch) quarter inch diameter hole. The invention virtually eliminated potential hot spots (high heat flux) forward, aft and around the injection sites and avoided the classic separation bubble that forms in front of a normal jet and recirculation eddies that form downstream of the jet. The elimination of these disturbances led to improved supersonic surface injection, penetration with uniform flow, and greater reliability.
A primary objective is to provide a slender flush injector which has better penetration, mixing and stable burning quickly without the high drag, intrusive, pressure wave, shock generating disturbances and resultant combustion instability defects present with state-of-the-art injectors and injector systems. A second purpose or objective of the invention was to substantially totally eliminate any not to be ignored intrusive/drag effects when the injector or injectors are not being used.
Another object of the instant invention was to achieve a wide range of supersonic flow injectors having uniform supersonic flow velocity through out the entire length of a non-intrusive jet with increased penetration into the cross flow and with better mixing over a large operational range of injectant pressures including injectant pressures which significantly exceed combustor, freestream (crossflow) total pressure.
Another object of the instant invention was the provision of a family or series of Mach number and mass flow specific supersonic slender injectors that have superior penetration and mixing capabilities in a wide range of supersonic crossflow conditions while substantially eliminating the intrusiveness generated flow disturbances, separation bubble, hot spots, pressure waves, shocks and recirculation eddies associated with the quarter inch hole injectors and the wedge configuration of the Bulman ""787 patent.
A better understanding of the invention will be had when reference is made to the following BRIEF DESCRIPTION OF THE DRAWINGS, DESCRIPTION OF THE INVENTION and CLAIMS which follow hereinbelow.